Gas turbine engine airfoil

ABSTRACT

An airfoil for a turbine engine includes pressure and suction sides that extend in a radial direction from a 0% span position at an inner flow path location to a 100% span position at an airfoil tip. The airfoil has a relationship between a stacking offset and a span position that is at least a third order polynomial curve that includes at least one positive and negative slope. The positive slope leans aftward and the negative slope leans forward relative to an engine axis. The positive slope crosses an initial axial stacking offset corresponding to the 0% span position at a zero-crossing position. A first axial stacking offset X 1  is provided from the zero-crossing position to a negative-most value on the curve. A second axial stacking offset X 2  is provided from the zero-crossing position to a positive-most value on the curve. A ratio of the second to first axial stacking offset X 2 /X 1  is less than 2.0.

CROSS-REFERENCE TO RELATED APPLICATIONS

This application is a Continuation of U.S. application Ser. No.14/625,413 filed Feb. 18, 2015, which claims priority to U.S.Provisional Application No. 61/941,781, which was filed on Feb. 19, 2014and is incorporated herein by reference.

BACKGROUND

This disclosure relates generally to an airfoil for gas turbine engines,and more particularly to a gas turbine engine fan or compressor blade,and a relationship between an axial stacking offset relative to span.

A turbine engine such as a gas turbine engine typically includes a fansection, a compressor section, a combustor section and a turbinesection. Air entering the compressor section is compressed and deliveredinto the combustor section where it is mixed with fuel and ignited togenerate a high-speed exhaust gas flow. The high-speed exhaust gas flowexpands through the turbine section to drive the compressor and the fansection. The compressor section typically includes low and high pressurecompressors, and the turbine section includes low and high pressureturbines.

The propulsive efficiency of a gas turbine engine depends on manydifferent factors, such as the design of the engine and the resultingperformance debits on the fan that propels the engine. As an example,the fan may rotate at a high rate of speed such that air passes over thefan airfoils at transonic or supersonic speeds. The fast-moving aircreates flow discontinuities or shocks that result in irreversiblepropulsive losses. Additionally, physical interaction between the fanand the air causes downstream turbulence and further losses. Althoughsome basic principles behind such losses are understood, identifying andchanging appropriate design factors to reduce such losses for a givenengine architecture has proven to be a complex and elusive task.

SUMMARY

In one exemplary embodiment, an airfoil for a turbine engine includespressure and suction sides that extend in a radial direction from a 0%span position at an inner flow path location to a 100% span position atan airfoil tip. The airfoil has a relationship between a stacking offsetand a span position that is at least a third order polynomial curve thatincludes at least one positive and negative slope. The positive slopeleans aftward and the negative slope leans forward relative to an engineaxis. The positive slope crosses an initial axial stacking offsetcorresponding to the 0% span position at a zero-crossing position. Afirst axial stacking offset X1 is provided from the zero-crossingposition to a negative-most value on the curve. A second axial stackingoffset X2 is provided from the zero-crossing position to a positive-mostvalue on the curve. A ratio of the second to first axial stacking offsetX2/X1 is less than 2.0.

In a further embodiment of the above, the curve has at least onecritical and one inflection point.

In a further embodiment of any of the above, the airfoil extends from aroot. A zero axial stacking offset corresponds to axial center of theroot.

In a further embodiment of any of the above, the curve has an initialnegative slope.

In a further embodiment of any of the above, a critical point is in therange of 25-50% span and provides the negative-most value.

In a further embodiment of any of the above, the critical point is inthe range of 35-45% span.

In a further embodiment of any of the above, the ratio of the second tofirst axial stacking offset increases as a span position for thezero-crossing position decreases.

In a further embodiment of any of the above, the zero-crossing point isin the range of 65-75% span position.

In a further embodiment of any of the above, the critical point has aRd/Xd in a range of −26 to −28.

In a further embodiment of any of the above, the zero-crossing point isin the range of 50-60% span position.

In a further embodiment of any of the above, the critical point has aRd/Xd in a range of −21 to −24.

In a further embodiment of any of the above, the zero-crossing point isin the range of 70-80% span position. The critical point has a Rd/Xd ina range of −33 to −37.

In a further embodiment of any of the above, a second critical point isthe positive-most value. The second critical point has a Rd/Xd in arange of 38 to 42.

In a further embodiment of any of the above, the positive slope extendsfrom the critical point to the inflection point at a first rate. Anending slope extends to the 100% span position. The ending slope has asecond rate that is less than the first rate or negative.

In a further embodiment of any of the above, the 100% span position onthe positive slope provides the positive-most value.

In a further embodiment of any of the above, the 100% span position hasa Rd/Xd in a range of 43 to 46.

In a further embodiment of any of the above, the 100% span position hasa Rd/Xd in a range of 58 to 62.

In a further embodiment of any of the above, the airfoil is in one of afan section, a compressor section, and a turbine section

In a further embodiment of any of the above, a ratio of the second tofirst axial stacking offset X2/X1 is less than one. A ratio of thesecond to first axial stacking offset X2/X1 is less than one half.

In another exemplary embodiment, a gas turbine engine includes acombustor section arranged between a compressor section and a turbinesection. A fan section has an array of twenty-six or fewer fan blades.The fan section has a low fan pressure ratio of less than 1.55. A gearedarchitecture couples the fan section to the turbine section or thecompressor section. The fan blades include an airfoil that has pressureand suction sides. The airfoil extends in a radial direction from a 0%span position at an inner flow path location to a 100% span position atan airfoil tip. The airfoil has a relationship between a stacking offsetand a span position that is at least a third order polynomial curve thatincludes at least one positive and negative slope. The positive slopeleans aftward and the negative slope leans forward relative to an engineaxis. The positive slope crosses an initial axial stacking offsetcorresponding to the 0% span position at a zero-crossing position. Afirst axial stacking offset X1 is provided from the zero-crossingposition to a negative-most value on the curve. A second axial stackingoffset X2 is provided from the zero-crossing position to a positive-mostvalue on the curve. A ratio of the second to first axial stacking offsetX2/X1 is less than 2.0.

BRIEF DESCRIPTION OF THE DRAWINGS

The disclosure can be further understood by reference to the followingdetailed description when considered in connection with the accompanyingdrawings wherein:

FIG. 1 schematically illustrates a gas turbine engine embodiment.

FIG. 2A is a perspective view of a portion of a fan section.

FIG. 2B is a schematic cross-sectional view of the fan section.

FIG. 2C is a cross-sectional view a fan blade taken along line 2C-2C inFIG. 2B.

FIG. 3A is a schematic view of fan blade span positions.

FIG. 3B is a schematic view of a cross-section of a fan blade section ata particular span position and its axial stacking offset.

FIG. 4A illustrates a relationship between an axial stacking offset anda span position for a set of first example airfoils.

FIG. 4B illustrates a relationship between an axial stacking offset anda span position for a set of second example airfoils.

FIG. 4C illustrates a relationship between an axial stacking offset anda span position for a set of third example airfoils.

The embodiments, examples and alternatives of the preceding paragraphs,the claims, or the following description and drawings, including any oftheir various aspects or respective individual features, may be takenindependently or in any combination. Features described in connectionwith one embodiment are applicable to all embodiments, unless suchfeatures are incompatible.

DETAILED DESCRIPTION

FIG. 1 schematically illustrates a gas turbine engine 20. The gasturbine engine 20 is disclosed herein as a two-spool turbofan thatgenerally incorporates a fan section 22, a compressor section 24, acombustor section 26 and a turbine section 28. Alternative engines mightinclude an augmenter section (not shown) among other systems orfeatures. The fan section 22 drives air along a bypass flow path B in abypass duct defined within a nacelle 15, while the compressor section 24drives air along a core flow path C for compression and communicationinto the combustor section 26 then expansion through the turbine section28. Although depicted as a two-spool turbofan gas turbine engine in thedisclosed non-limiting embodiment, it should be understood that theconcepts described herein are not limited to use with two-spoolturbofans as the teachings may be applied to other types of turbineengines including three-spool architectures. That is, the disclosedairfoils may be used for engine configurations such as, for example,direct fan drives, or two- or three-spool engines with a speed changemechanism coupling the fan with a compressor or a turbine sections.

The exemplary engine 20 generally includes a low speed spool 30 and ahigh speed spool 32 mounted for rotation about an engine centrallongitudinal axis X relative to an engine static structure 36 viaseveral bearing systems 38. It should be understood that various bearingsystems 38 at various locations may alternatively or additionally beprovided, and the location of bearing systems 38 may be varied asappropriate to the application.

The low speed spool 30 generally includes an inner shaft 40 thatinterconnects a fan 42, a first (or low) pressure compressor 44 and afirst (or low) pressure turbine 46. The inner shaft 40 is connected tothe fan 42 through a speed change mechanism, which in exemplary gasturbine engine 20 is illustrated as a geared architecture 48 to drivethe fan 42 at a lower speed than the low speed spool 30. The high speedspool 32 includes an outer shaft 50 that interconnects a second (orhigh) pressure compressor 52 and a second (or high) pressure turbine 54.A combustor 56 is arranged in exemplary gas turbine 20 between the highpressure compressor 52 and the high pressure turbine 54. A mid-turbineframe 57 of the engine static structure 36 is arranged generally betweenthe high pressure turbine 54 and the low pressure turbine 46. Themid-turbine frame 57 further supports bearing systems 38 in the turbinesection 28. The inner shaft 40 and the outer shaft 50 are concentric androtate via bearing systems 38 about the engine central longitudinal axisX which is collinear with their longitudinal axes.

The core airflow is compressed by the low pressure compressor 44 thenthe high pressure compressor 52, mixed and burned with fuel in thecombustor 56, then expanded over the high pressure turbine 54 and lowpressure turbine 46. The mid-turbine frame 57 includes airfoils 59 whichare in the core airflow path C. The turbines 46, 54 rotationally drivethe respective low speed spool 30 and high speed spool 32 in response tothe expansion. It will be appreciated that each of the positions of thefan section 22, compressor section 24, combustor section 26, turbinesection 28, and fan drive gear system 48 may be varied. For example,gear system 48 may be located aft of combustor section 26 or even aft ofturbine section 28, and fan section 22 may be positioned forward or aftof the location of gear system 48.

The engine 20 in one example is a high-bypass geared aircraft engine. Ina further example, the engine 20 bypass ratio is greater than about six(6), with an example embodiment being greater than about ten (10), thegeared architecture 48 is an epicyclic gear train, such as a planetarygear system or other gear system, with a gear reduction ratio of greaterthan about 2.3 and the low pressure turbine 46 has a pressure ratio thatis greater than about five. In one disclosed embodiment, the engine 20bypass ratio is greater than about ten (10:1), the fan diameter issignificantly larger than that of the low pressure compressor 44, andthe low pressure turbine 46 has a pressure ratio that is greater thanabout five (5:1). Low pressure turbine 46 pressure ratio is pressuremeasured prior to inlet of low pressure turbine 46 as related to thepressure at the outlet of the low pressure turbine 46 prior to anexhaust nozzle. The geared architecture 48 may be an epicyclic geartrain, such as a planetary gear system or other gear system, with a gearreduction ratio of greater than about 2.3:1. It should be understood,however, that the above parameters are only exemplary of one embodimentof a geared architecture engine and that the present invention isapplicable to other gas turbine engines including direct driveturbofans.

The example gas turbine engine includes the fan 42 that comprises in onenon-limiting embodiment less than about twenty-six (26) fan blades. Inanother non-limiting embodiment, the fan section 22 includes less thanabout twenty (20) fan blades. Moreover, in one disclosed embodiment thelow pressure turbine 46 includes no more than about six (6) turbinerotors schematically indicated at 34. In another non-limiting exampleembodiment the low pressure turbine 46 includes about three (3) turbinerotors. A ratio between the number of fan blades 42 and the number oflow pressure turbine rotors is between about 3.3 and about 8.6. Theexample low pressure turbine 46 provides the driving power to rotate thefan section 22 and therefore the relationship between the number ofturbine rotors 34 in the low pressure turbine 46 and the number ofblades 42 in the fan section 22 disclose an example gas turbine engine20 with increased power transfer efficiency.

A significant amount of thrust is provided by the bypass flow B due tothe high bypass ratio. The fan section 22 of the engine 20 is designedfor a particular flight condition—typically cruise at about 0.8 Mach andabout 35,000 feet. The flight condition of 0.8 Mach and 35,000 ft, withthe engine at its best fuel consumption—also known as “bucket cruiseThrust Specific Fuel Consumption (‘TSFC’)”—is the industry standardparameter of lbm of fuel being burned divided by lbf of thrust theengine produces at that minimum point. “Low fan pressure ratio” is thepressure ratio across the fan blade alone, without a Fan Exit Guide Vane(“FEGV”) system. The low fan pressure ratio as disclosed hereinaccording to one non-limiting embodiment is less than about 1.55. Inanother non-limiting embodiment the low fan pressure ratio is less thanabout 1.52. In another non-limiting embodiment the low fan pressureratio is less than about 1.50. In another non-limiting embodiment thelow fan pressure ratio is less than about 1.48. In another non-limitingembodiment the low fan pressure ratio is less than about 1.46. Inanother non-limiting embodiment the low fan pressure ratio is less thanabout 1.44. In another non-limiting embodiment the low fan pressureratio is from 1.1 to 1.45. “Low corrected fan tip speed” is the actualfan tip speed in ft/sec divided by an industry standard temperaturecorrection of [(Tram ° R)/(518.7° R)]^(0.5). The “Low corrected fan tipspeed” as disclosed herein according to one non-limiting embodiment isless than about 1200 ft/second.

Referring to FIG. 2A-2C, the fan blade 42 is supported by a fan hub 60that is rotatable about the axis X. Each fan blade 42 includes anairfoil 64 extending in a radial span direction R from a root 62 to atip 66. A 0% span position corresponds to a section of the airfoil 64 atthe inner flow path (e.g., a platform), and a 100% span positioncorresponds to a section of the airfoil 64 at the tip 66.

The root 62 is received in a correspondingly shaped slot in the fan hub60. The airfoil 64 extends radially outward of the platform, whichprovides the inner flow path. The platform may be integral with the fanblade or separately secured to the fan hub, for example. A spinner 66 issupported relative to the fan hub 60 to provide an aerodynamic innerflow path into the fan section 22.

The airfoil 64 has an exterior surface 76 providing a contour thatextends from a leading edge 68 aftward in a chord-wise direction H to atrailing edge 70, as shown in FIG. 2C. Pressure and suction sides 72, 74join one another at the leading and trailing edges 68, 70 and are spacedapart from one another in an airfoil thickness direction T. An array ofthe fan blades 42 are positioned about the axis X in a circumferentialor tangential direction Y. Any suitable number of fan blades may be usedin a given application.

The exterior surface 76 of the airfoil 64 generates lift based upon itsgeometry and directs flow along the core flow path C. The fan blade 42may be constructed from a composite material, or an aluminum alloy ortitanium alloy, or a combination of one or more of these.Abrasion-resistant coatings or other protective coatings may be appliedto the fan blade 42.

One characteristic of fan blade performance relates to the fan blade'saxial stacking offset (X direction) relative to a particular spanposition (R direction). Referring to FIG. 3A, span positions areschematically illustrated from 0% to 100% in 10% increments. Eachsection at a given span position is provided by a conical cut thatcorresponds to the shape of the core flow path, as shown by the largedashed lines. In the case of a fan blade with an integral platform, the0% span position corresponds to the radially innermost location wherethe airfoil meets the fillet joining the airfoil to the platform. In thecase of a fan blade without an integral platform, the 0% span positioncorresponds to the radially innermost location where the discreteplatform meets the exterior surface of the airfoil. In addition tovarying with span, axial stacking offset varies between a hot, runningcondition and a cold, static (“on the bench”) condition.

The X_(CG) corresponds to the location of the center of gravity for aparticular section at a given span location relative to a referencepoint 80 in the X direction, as shown in FIG. 3B. The center of gravityassumes a homogenous material. The reference point 80 is the axialcenter of the root, and X_(d) corresponds to the axial distance from thereference point 80 to the center of gravity. A positive X is on the aftside of reference point 80, and a negative X is on the forward side ofthe reference point 80. A positive slope is where the leading edge leansaftward, and a negative slope is where the leading edge leans forwardrelative to the engine's axis X. The inflection points are indicated byan “x” on the curve.

The axial stacking offset X_(d) may be expressed as an axial stackingoffset ratio R_(d)/X_(d), which is the ratio of the radial distanceR_(d) to a given span position, where R_(d) is the radial distance fromhub's rotational axis X to the tip of the leading edge 68, divided bythe axial distance X_(d). R_(d) as disclosed herein according to onenon-limiting embodiment is about 35-37 inches (0.89-0.94 meters). Inanother non-limiting embodiment R_(d) is about 27-29 inches (0.69-0.74meters). In another non-limiting embodiment R_(d) is about 39-41 inches(0.99-1.04 meters). One example prior art airfoil has an R_(d) of about57-59 inches (1.45-1.50 meters).

Example relationships between the axial stacking offset and the spanposition (% AVE SPAN), which is the average of the radial position atthe leading and trailing edges 68, 70, are shown in FIGS. 4A-4C forseveral example fan blades, each represented by a curve. Only one curvein each graph is discussed for simplicity. Each relationship is at leasta third order polynomial curve that includes at least one positive andnegative slope. In one example, the curve has at least one inflectionand one critical point. The curve initially has a negative slope andthen transitions to a positive slope at a critical point in the range of25-50% span. In the examples, the critical point is in the range ofabout 35-45% span. The critical point provides a negative-most value.The positive slope crosses an initial axial stacking offsetcorresponding to the 0% span position at a zero-crossing position (92 inFIG. 4A; 98 in FIG. 4B; 104 in FIG. 4C).

A first axial stacking offset is provided from the zero-crossingposition to the negative-most value on the curve (X1). A second axialstacking offset is provided from the zero-cros sing position to apositive-most value on the curve (X2). The ratio of the second to firstaxial stacking offset X2/X1 increases as a span position for thezero-crossing position decreases. A ratio of the second to first axialstacking offset X2/X1 is between 1.5 and 2.0 or less than 1.4, whichprovides a less dramatic aftward-facing hook at the tip of the bladethan some prior art geometries. In another example the ratio is lessthan one, and in another example the ratio is less than one half.Example prior art stacking offset ratios are 1.44 or greater than 2.0.

Referring to FIG. 4A, the curve has an initial negative slope to acritical point 90 in the range of 35-45% span, and in one example,around 40% span. The critical point 90 has an R_(d)/X_(d) in the rangeof −26 to −28. The positive slope extends from the critical point 90 toan inflection point 94 at a first rate. An ending slope extends to the100% span position from the inflection point 94. The ending slope havinga second rate that is less than the first rate. The positive slopeincludes a zero-crossing point 92 in the range of 65-75% span position.The ratio of the second to first axial stacking offset X2 _(A)/X1 _(A)is around 0.6. At the 100% span location, the axial stacking offsetratio R_(d)/X_(d) is in the range of 43 to 46. The 100% span position onthe positive slope provides the positive-most value.

Referring to FIG. 4B, the curve has an initial negative slope to acritical point 96 in the range of 30-40% span, and in one example,around 35% span. The critical point 96 has an R_(d)/X_(d) in the rangeof −21 to −24. The positive slope extends from the critical point 96 toan inflection point 100 at a first rate. An ending slope extends to the100% span position from the second inflection point 100. The endingslope having a second rate that is less than the first rate. Thepositive slope includes a zero-crossing point 98 in the range of 75-85%span position. The ratio of the second to first axial stacking offset X2_(B)/X1 _(B) is around 0.4. At the 100% span location, the axialstacking offset ratio R_(d)/X_(d) is in the range of 58 to 62. The 100%span position on the positive slope provides the positive-most value.

Referring to FIG. 4C, the curve has an initial negative slope to a firstcritical point 102 in the range of 30-40% span, and in one example,around 35% span. The first critical point 102 has an R_(d)/X_(d) in therange of −33 to −37. The positive slope extends from the first criticalpoint 102 to an inflection point 106 and a second critical point 108that provides the positive-most value. An ending slope extends to the100% span position from the second critical point 108. The ending slopehaving a second rate that is negative. The positive slope includes azero-crossing point 104 in the range of 55-65% span position. The ratioof the second to first axial stacking offset X2 _(C)/X1 _(C) is around1.5. At the 100% span location, the axial stacking offset ratioR_(d)/X_(d) is in the range of 38 to 42.

The axial stacking offset is a is a hot, running condition along thespan of the airfoils 64 relate to the contour of the airfoil and providenecessary fan operation in cruise at the lower, preferential speedsenabled by the geared architecture 48 in order to enhance aerodynamicfunctionality and thermal efficiency. As used herein, the hot, runningcondition is the condition during cruise of the gas turbine engine 20.For example, the axial stacking offsets in the hot, running conditioncan be determined in a known manner using numerical analysis, such asfinite element analysis.

It should also be understood that although a particular componentarrangement is disclosed in the illustrated embodiment, otherarrangements will benefit herefrom. For example, the disclosed airfoilmay be used in any one of a fan section, a compressor section, and aturbine section. Although particular step sequences are shown,described, and claimed, it should be understood that steps may beperformed in any order, separated or combined unless otherwise indicatedand will still benefit from the present invention.

Although the different examples have specific components shown in theillustrations, embodiments of this invention are not limited to thoseparticular combinations. It is possible to use some of the components orfeatures from one of the examples in combination with features orcomponents from another one of the examples.

Although an example embodiment has been disclosed, a worker of ordinaryskill in this art would recognize that certain modifications would comewithin the scope of the claims. For that reason, the following claimsshould be studied to determine their true scope and content.

What is claimed is:
 1. An airfoil for a turbine engine comprising:pressure and suction sides extending in a radial direction from a 0%span position at an inner flow path location to a 100% span position atan airfoil tip, wherein the airfoil has a relationship between astacking offset and a span position that is at least a third orderpolynomial curve that includes at least one positive and negative slope,the positive slope leans aftward and the negative slope leans forwardrelative to an engine axis, wherein the positive slope crosses aninitial axial stacking offset corresponding to the 0% span position at azero-crossing position, a first axial stacking offset X1 is providedfrom the zero-crossing position to a negative-most value on the curve, asecond axial stacking offset X2 is provided from the zero-crossingposition to a positive-most value on the curve, a ratio of the second tofirst axial stacking offset X2/X1 is less than 2.0.
 2. The airfoilaccording to claim 1, wherein the curve has at least one critical andone inflection point.
 3. The airfoil according to claim 2, wherein theairfoil extends from a root, and a zero axial stacking offsetcorresponds to axial center of the root.
 4. The airfoil according toclaim 2, wherein the curve has an initial negative slope.
 5. The airfoilaccording to claim 4, wherein a critical point is in the range of 25-50%span and provides the negative-most value.
 6. The airfoil according toclaim 5, wherein the critical point is in the range of 35-45% span. 7.The airfoil according to claim 5, wherein the ratio of the second tofirst axial stacking offset increases as a span position for thezero-crossing position decreases.
 8. The airfoil according to claim 7,wherein the zero-crossing point is in the range of 65-75% span position.9. The airfoil according to claim 8, wherein the critical point has aR_(d)/X_(d) in a range of −26 to −28.
 10. The airfoil according to claim7, wherein the zero-crossing point is in the range of 50-60% spanposition.
 11. The airfoil according to claim 10, wherein the criticalpoint has a R_(d)/X_(d) in a range of −21 to −24.
 12. The airfoilaccording to claim 7, wherein the zero-crossing point is in the range of70-80% span position, and wherein the critical point has a R_(d)/X_(d)in a range of −33 to −37.
 13. The airfoil according to claim 5, whereina second critical point is the positive-most value, and wherein thesecond critical point has a R_(d)/X_(d) in a range of 38 to
 42. 14. Theairfoil according to claim 5, wherein the positive slope extends fromthe critical point to the inflection point at a first rate, an endingslope extends to the 100% span position, the ending slope having asecond rate that is less than the first rate or negative.
 15. Theairfoil according to claim 5, wherein the 100% span position on thepositive slope provides the positive-most value.
 16. The airfoilaccording to claim 14, wherein the 100% span position has a R_(d)/X_(d)in a range of 43 to
 46. 17. The airfoil according to claim 14, whereinthe 100% span position has a R_(d)/X_(d) in a range of 58 to
 62. 18. Theairfoil according to claim 1, wherein the airfoil is in one of a fansection, a compressor section, and a turbine section
 19. The airfoilaccording to claim 1, wherein a ratio of the second to first axialstacking offset X2/X1 is less than one, and wherein a ratio of thesecond to first axial stacking offset X2/X1 is less than one half.
 20. Agas turbine engine comprising: a combustor section arranged between acompressor section and a turbine section; a fan section having an arrayof twenty-six or fewer fan blades, wherein the fan section has a low fanpressure ratio of less than 1.55; a geared architecture coupling the fansection to the turbine section or the compressor section; and whereinthe fan blades include an airfoil having pressure and suction sides, theairfoil extends in a radial direction from a 0% span position at aninner flow path location to a 100% span position at an airfoil tip,wherein the airfoil has a relationship between a stacking offset and aspan position that is at least a third order polynomial curve thatincludes at least one positive and negative slope, the positive slopeleans aftward and the negative slope leans forward relative to an engineaxis, wherein the positive slope crosses an initial axial stackingoffset corresponding to the 0% span position at a zero-crossingposition, a first axial stacking offset X1 is provided from thezero-crossing position to a negative-most value on the curve, a secondaxial stacking offset X2 is provided from the zero-crossing position toa positive-most value on the curve, a ratio of the second to first axialstacking offset X2/X1 is less than 2.0.